Turbine guide vane and cooling structure thereof

Abstract

The turbine guide vane cooling structure is arranged on a turbine guide vane, the turbine guide vane comprises a suction surface side vane body wall body and a pressure surface side vane body wall body, and further comprises a front cavity and a rear cavity which are divided by a cavity division partition plate located in a vane body, and further comprises a partition plate with the two sides connected with the cavity division partition plate and the suction surface side vane body wall body respectively, the partition plate is used for dividing the front cavity into an impact cavity, the impact cavity is located in the blade back cheek area, a partition plate impact hole is formed in the partition plate and used for communicating the impact cavity with the front cavity, and the cavity dividing partition plate is provided with a cavity dividing impact hole used for communicating the impact cavity with the rear cavity; wherein the front cavity is provided with a single-layer wall and an impact lining, and the rear cavity is provided with a double-layer wall cooling structure. The invention further provides the turbine guide vane with the cooling structure. The cooling structure has a better cooling effect.

Description

Technical Field

[0001] This invention relates to the field of turbine blades, and more specifically to the field of cooling structures. Background Technology

[0002] Currently, the inlet temperature of high-pressure turbines in aero engines is close to 2000K. With the development of modern aero engine technology, the turbine inlet temperature continues to rise and will reach the level of 2100K to 2200K.

[0003] The basic cooling methods for high-pressure turbine blades in modern aero-engines include impingement, turbulence, and film cooling. Turbine blades often employ a double-wall cooling structure. The double wall improves heat transfer in the solid portion of the blade surface. However, when applying irregularly shaped film cooling holes to the outer wall of the double-wall design, in addition to the manufacturing difficulties, the thin wall also limits the expansion area of ​​the irregularly shaped film cooling holes at the blade's outer wall outlet, resulting in a relatively limited improvement in film cooling efficiency compared to round holes.

[0004] How to further improve the cooling effect of turbine guide vanes is an urgent problem to be solved. Summary of the Invention

[0005] One object of the present invention is to provide a turbine guide vane cooling structure with better cooling effect.

[0006] To achieve the above objectives, a turbine guide vane cooling structure is provided on the turbine guide vane. The turbine guide vane includes a suction-side blade wall and a pressure-side blade wall, and also includes a front cavity and a rear cavity divided by a partition plate located inside the blade body. It further includes partition plates connected to the partition plate and the suction-side blade wall on both sides, respectively, for dividing the front cavity into an impact cavity. The impact cavity is located in the blade back cheek region. The partition plate has partition impact holes for connecting the impact cavity and the front cavity. The partition plate also has partition impact holes for connecting the impact cavity and the rear cavity. The front cavity has a single-layer wall and an impact bushing, while the rear cavity has a double-layer wall cooling structure.

[0007] In one or more embodiments, the impact bushing is provided with bushing impact holes, and the single-layer wall is provided with air film holes.

[0008] In one or more embodiments, at least some of the air film pores on the single-layer wall are irregularly shaped air film pores.

[0009] In one or more embodiments, the double-wall cooling structure includes a suction-side cooling structure and a pressure-side cooling structure.

[0010] In one or more embodiments, the pressure-side cooling structure includes an inner wall, an outer wall, and a gap turbulence column located between them, wherein the inner wall and the outer wall are provided with film cooling holes, or the structure includes an inner wall, an outer wall, and film cooling holes located on the inner wall and the outer wall; the suction-side cooling structure includes an inner wall, an outer wall, and a gap turbulence column located between them, wherein the inner wall and the outer wall are provided with film cooling holes, or the structure includes an inner wall, an outer wall, and film cooling holes located on the inner wall and the outer wall.

[0011] In one or more embodiments, the air film pores on the inner wall and the outer wall are one or more of the following: conical air film pores, scoop-shaped air film pores, expanded air film pores, and W-shaped air film pores.

[0012] In one or more embodiments, the turbine guide vane cooling structure further includes a tail section partition, the two sides of which are connected to the suction-side blade wall and the pressure-side blade wall, thereby dividing the rear cavity into a trailing edge cavity.

[0013] In one or more embodiments, the double-wall cooling structure includes an inner wall and an outer wall, with the tail ends of the inner wall located on the pressure side and the suction side connected to form the tail partition.

[0014] In one or more embodiments, the trailing edge cavity is provided with multiple rows of trailing edge baffles and trailing edge slits.

[0015] In one or more embodiments, the diameter of the trailing edge spoiler column ranges from 1 mm to 1.8 mm.

[0016] In one or more embodiments, the wall thickness of the single-layer wall ranges from 1.6 mm to 2.2 mm, and / or the wall thickness of the impact bushing ranges from 0.2 mm to 0.5 mm, and / or the impact distance between the impact bushing and the single-layer wall ranges from 0.4 mm to 1.0 mm.

[0017] In one or more embodiments, the wall thickness of the double-walled structure ranges from 2.1 mm to 2.7 mm, and / or

[0018] The diameter of the air film pores on the inner wall ranges from 0.8 mm to 1.0 mm, and / or the diameter of the gap turbulence column ranges from 0.8 mm to 1.4 mm.

[0019] In one or more embodiments, the diameter of the cavity impact hole ranges from 1.2 mm to 1.4 mm.

[0020] In one or more embodiments, the suction-side blade wall forming the impact cavity is provided with multiple exhaust film holes, with the outlet of the last exhaust film hole located on the front side of the blade throat.

[0021] In one or more embodiments, the turbine guide vane cooling structure further includes a rear chamber air inlet and a front chamber air inlet, wherein the rear chamber air inlet is located at the upper end of the turbine guide vane and communicates with the rear chamber, and the front chamber air inlet is located at the upper end of the turbine guide vane and communicates with the front chamber.

[0022] Another object of the present invention is to provide a turbine guide vane including the turbine guide vane cooling structure described above.

[0023] The aforementioned turbine guide vane cooling structure, based on fully considering the external flow heat transfer characteristics at different locations of the blade, leverages the advantages of single-layer wall, double-layer wall, and irregularly shaped air film pore structures in different regions. Furthermore, an impact chamber is added in the blade back and cheek area to further cool the blade back and cheek area, thereby improving the overall air film coverage effect from the blade back and cheek area to the trailing edge area, resulting in a significant improvement in cooling performance. Attached Figure Description

[0024] The above and other features, properties and advantages of the present invention will become more apparent from the following description taken in conjunction with the accompanying drawings and embodiments, wherein:

[0025] Figure 1 This is a schematic diagram of the blade basin of a turbine guide vane;

[0026] Figure 2 This is a schematic diagram of the back surface of a turbine guide vane;

[0027] Figure 3 yes Figure 2 Cross-sectional view along the AA direction;

[0028] Figure 4 yes Figure 2 Cross-sectional view along the BB direction;

[0029] Figure 5 yes Figure 2 Cross-sectional view along the CC direction;

[0030] Figures 6A-6B This is a schematic diagram of a conical air film orifice;

[0031] Figures 7A-7B This is a schematic diagram of a sieve-shaped air film vent.

[0032] Figure 8 This is a structural schematic diagram of an impact bushing;

[0033] Figure 9 This is a cross-sectional schematic diagram of one embodiment of a double-walled cooling unit;

[0034] Figure 10 yes Figure 9 Cross-sectional view along the DD direction;

[0035] Figure 11 This is a schematic diagram of the airflow path within the front cavity channel along the blade spanwise direction.

[0036] Figure 12 This is a schematic diagram of the airflow path within the rear cavity flow channel;

[0037] Figure 13A This is a simulation diagram of the air film cooling effect of a structure combining double-walled and circular air film pores.

[0038] Figure 13B It is a simulation diagram of the air film cooling effect of a single-layer wall structure with irregularly shaped air film pores;

[0039] Figures 14A-14B This is a schematic diagram of an expanded film pore;

[0040] Figures 15A-15B This is a schematic diagram of a W-shaped air film vent;

[0041] Figure 16 This is a schematic diagram of another embodiment of a turbine guide vane;

[0042] Figure 17 This is a cross-sectional schematic diagram of another embodiment of the double-walled cooling unit;

[0043] Figure 18 yes Figure 17 Cross-sectional view along the EE direction. Detailed Implementation

[0044] The present invention will be further described below with reference to specific embodiments and accompanying drawings. More details are set forth in the following description in order to provide a full understanding of the present invention. However, the present invention can obviously be implemented in many other ways different from those described herein. Those skilled in the art can make similar extensions and derivations based on actual application situations without departing from the spirit of the present invention. Therefore, the scope of protection of the present invention should not be limited by the content of this specific embodiment.

[0045] It should be noted that these and other accompanying drawings are merely examples and are not drawn to scale, and should not be construed as limiting the scope of protection of the present invention.

[0046] Figure 1 and Figure 2 The blade face and back face of turbine guide vane 1 are shown respectively. Figure 3 The cross-section of turbine guide vane 1 along the AA direction is shown. Figure 4 and Figure 5 Show along Figure 3 Cross-sectional views in the BB and CC directions.

[0047] The turbine guide vane cooling structure disclosed herein is disposed on the turbine guide vane 1. The turbine guide vane 1 includes a blade body 7 and a rim plate 8. The blade body 7 includes a leading edge 101, a trailing edge 104, a pressure surface 102 (blade basin surface), and a suction surface 103 (blade back surface). The wall forming the blade body 7 includes a suction surface side blade body wall 701 and a pressure surface side blade body wall 702. It also includes a front cavity I and a rear cavity II divided by a partition plate 11 located inside the blade body. The front cavity I and the rear cavity II are supplied with cooling air by two separate channels.

[0048] The cooling structure also includes a baffle 5, which is connected to a partition baffle 11 and a suction-side blade wall 701 on both sides, respectively, to divide the front cavity I into an impact cavity III. The impact cavity III is located in the blade back cheek region R, that is, on the back of the front cavity I near the blade throat Y. The "throat" is the area with the smallest cross-sectional area between two circumferentially adjacent turbine guide vanes 1. In one embodiment, the turbine guide vane cooling structure also includes a tail baffle 6, which is connected to the suction-side blade wall 701 and the pressure-side blade wall 702 on both sides, thereby dividing the rear cavity II into a trailing edge cavity IV.

[0049] The partition 5 is provided with a partition impact hole 51 for connecting the impact chamber III and the front chamber I. The partition 11 is provided with a partition impact hole 22 for connecting the impact chamber III and the rear chamber II.

[0050] The different chambers inside the turbine guide vane are described below.

[0051] The front chamber I is equipped with a single-layer wall 12 and an impact bushing 2. The single-layer wall 12 has a wall thickness of 1.6mm to 2.2mm, and the impact bushing 2 is arranged inside it. The area between the impact bushing 2 and the single-layer wall 12 is the impact flow region. One end of the impact flow region is connected to the front chamber I through the bushing impact hole 21 on the impact bushing 2, and the other end is connected to the outside of the blade through the film cooling holes on the surface of the single-layer wall. The front chamber I as a whole adopts a cooling design structure that combines a single-layer wall and irregularly shaped film cooling holes.

[0052] The single-layer wall 12 of the front cavity I is provided with air film pores, preferably at least part of the air film pores on the single-layer wall are irregular air film pores. Irregular air film pores include, but are not limited to, one or more of the following: scoop-shaped air film pores, expansion-shaped air film pores, and W-shaped air film pores.

[0053] Furthermore, the front end of the single-layer wall 12 is a leading edge 101, which preferably adopts a conical air film orifice 111. Since the leading edge 101 is relatively wide, the number of air film orifices here can be set to 8 rows. (See the conical air film orifice 111 for reference.) Figures 6A-6B As shown, the aperture c1 is approximately 0.5 mm to 0.6 mm, the yaw angle α1 is approximately 50° to 60°, and the expansion angle β1 is approximately 8° to 12°. The ratio of the length s1 of the expansion section 113 to the length l1 of the straight hole section 112 is (1.2 to 1.9):1.

[0054] The pressure surface 102 and suction surface 103 of the front chamber I preferably adopt a scoop-shaped air film orifice 121. (See reference 121 for details.) Figures 7A-7B As shown, the aperture c2 is approximately 0.5 mm to 0.7 mm, the tilt angle θ2 is approximately 30° to 50°, the back tilt angle ψ2 is approximately 9° to 12°, the expansion angle β2 is approximately 10° to 13°, and the ratio of the length l2 of the straight hole section 122 to the length s2 of the expansion section 123 is (1.0 to 1.5): 1.

[0055] like Figure 8 As shown, the wall thickness δ1 of the impact bushing 2 is 0.2mm to 0.5mm, and the impact bushing 2 is provided with bushing impact holes 21. The hole diameter of the bushing impact holes 21 is preferably 0.4mm to 0.8mm. The impact distance between the impact bushing 2 and the inner wall of the single-layer wall 12 is 0.4mm to 1.0mm.

[0056] Because the front cavity I has a dense array of film cooling holes, the wide coverage area and good effect of the irregularly shaped film cooling holes can provide more efficient film protection for the blade. The use of the impact bushing 2 can enhance the impact heat transfer on the inner wall surface. Although the heat transfer of the solid part of the blade wall is weaker than that of the double-walled blade, the single-walled blade has a thicker wall. In particular, the use of irregularly shaped film cooling holes, such as one or more of the following types, such as scoop-shaped film cooling holes, expanded film cooling holes, and W-shaped film cooling holes, can effectively expand the outlet section of the film cooling holes and increase the outlet area. Therefore, when applying irregularly shaped film cooling holes on a single-walled blade, the outlet velocity of the cold air can be significantly reduced, and the characteristics of the cold air at the outlet of the film cooling hole, such as being difficult to separate and having a long coverage area, can be fully utilized.

[0057] The turbine guide vane cooling structure described in this disclosure further cools the turbine blade's gill region, thereby significantly improving the overall film cooling efficiency of the blade. Based on this, furthermore, an impact chamber III is created within the front chamber I, located on the back surface of the front chamber blade near the blade's "throat." Impact chamber III is referenced... Figures 3 to 5 It is understood that the impact chamber III is connected to the front chamber I and the rear chamber II through the diaphragm impact hole 51 and the chamber-splitting impact hole 22, respectively. A portion of the cold air from the front chamber I and the rear chamber II enters the impact chamber III through the diaphragm impact hole 51 on the diaphragm 5 and the chamber-splitting impact hole 22 on the chamber-splitting diaphragm 11, respectively, and the impact improves the heat exchange effect of the inner wall of the blade back.

[0058] The diameter range of the cavity impact hole 22 is preferably 0.8 mm to 1.1 mm. The diameter range of the partition impact hole 51 is preferably 0.8 mm to 1.1 mm.

[0059] To improve the air film coverage effect, the air film holes 52 on the outer side of the impact chamber III are scoop-shaped air film holes. Furthermore, to reduce aerodynamic losses, the outlets of the last row of air film holes 52 on the suction surface 103 are located on the front side of the throat Y, that is, near the leading edge 101 of the throat Y. Figure 2 and Figure 3 As shown.

[0060] Thus, on the one hand, the impact cooling air from the front chamber I is provided through the diaphragm impact hole 51, which improves the internal cooling effect of the blade on the inner side of the blade back; on the other hand, the cooling air entering the rear chamber II increases the amount of cooling air entering the impact chamber III, which improves the cooling coverage effect of the irregular air film holes on the surface of the impact chamber III.

[0061] Figure 3 , Figures 9 to 10 The cross-sectional structure of rear cavity II is shown. The double-walled cooling structure of rear cavity II includes a suction-side cooling structure T and a pressure-side cooling structure W.

[0062] The pressure-side cooling structure includes an inner wall, an outer wall, and a gap turbulence column located between them. The inner wall and the outer wall are provided with film pores, or the structure includes an inner wall, an outer wall, and film pores located on the inner wall and the outer wall. The suction-side cooling structure includes an inner wall, an outer wall, and a gap turbulence column located between them. The inner wall and the outer wall are provided with film pores, or the structure includes an inner wall, an outer wall, and film pores located on the inner wall and the outer wall.

[0063] like Figure 3 As shown, in some embodiments, the pressure-side cooling structure includes an inner wall 14, an outer wall 13, and a gap turbulence column 18 located between them. The inner wall 14 and the outer wall 13 are provided with air film holes, forming a combination of impact, turbulence column, and air film on the pressure surface 102. The suction-side cooling structure includes an inner wall, an outer wall, and air film holes located on the inner wall and the outer wall, and is configured in a combination of impact and air film holes on the suction surface 103.

[0064] For example Figure 16 As shown, in some other embodiments, both the pressure-side cooling structure and the suction-side cooling structure include an inner wall, an outer wall, and air film pores located on the inner wall and the outer wall, forming a combination of impact and air film.

[0065] The width and number of channels in the double-walled cooling structure unit are determined based on the actual blade size, heat load distribution, and cooling air flow distribution. The film cooling orifice can be one or more of the following types: conical, scoop-shaped, expanded, or W-shaped.

[0066] Specifically, the double-walled cooling unit, which combines impact, turbulence columns, and film cooling, includes an outer wall 13, an inner wall 14, a double-walled partition 15, gap turbulence columns 18, and the resulting double-walled cavity 16. Cooling gas enters the double-walled cavity 16 from the rear cavity II through the inner wall impact holes 17, is turbulent by the gap turbulence columns 18, and then flows out of the blades through the circular film cooling holes 131. This structure forms an impact and turbulence heat transfer zone. One end of this zone is connected to the rear cavity II through the impact holes on the inner wall of the double-walled system, and the other end is connected to the outside through the film cooling holes on the outer wall of the double-walled system. Because the rear cavity II has a small number of film cooling holes, the double-walled structure reduces the wall surface heat transfer resistance and effectively improves the overall cooling effect.

[0067] In some embodiments, such as Figure 9 As shown, the ratio of the impact distance p2 of the double-walled channel to the diameter d1 of the inner wall impact hole 17 is (0.8~0.9):1, and the ratio of the diameter c3 of the circular air film hole 131 to the diameter d1 of the inner wall impact hole 17 is (0.5~0.9):1. Preferably, the impact hole is... Figure 10 The forked arrangement shown has gap turbulence columns 18 arranged between the rows of inner wall impact holes 17. The ratio of the hole spacing m1 to the hole diameter d1 of the inner wall impact holes 17 is (3~5):1, the ratio of the row spacing n2 to the hole diameter d1 is (3~6):1, and the ratio of the column spacing m2 to the column diameter e1 of the gap turbulence columns 18 is (2.5~4.5):1.

[0068] In other embodiments, the pressure surface 102 is configured as follows: Figure 17 and Figure 18 The impact and air film combination shown has the following characteristics: the ratio of the impact distance p2 of the double-walled channel to the diameter d2 of the inner wall impact hole 17 is (0.8-0.9):1; the ratio of the diameter c3 of the circular air film hole 131 to the diameter d2 of the inner wall impact hole 17 is (0.55-0.8):1. The inner wall impact holes 17 are arranged in a staggered pattern, with the ratio of the hole spacing m3 to the diameter d2 being (3.5-5.5):1, and the ratio of the row spacing n2 to the diameter d2 being (2.0-2.7):1.

[0069] For the blade tail, the turbine guide vane cooling structure also includes a trailing edge cavity IV formed by the tail partition 6. In some embodiments, the tail ends of the inner wall located on the pressure side and the suction side are connected to form the tail partition 6, such as... Figure 3 As shown.

[0070] The trailing edge cavity IV has a single-layer wall structure, with multiple rows of trailing edge baffles 19 inside the single-layer wall, and also a trailing edge slit 161, such as... Figure 1 and Figure 3 As shown. The diameter of the trailing edge spoiler column 19 ranges from 1mm to 1.8mm, and the specific number of rows and columns can be determined according to the actual blade size and airflow distribution.

[0071] Figure 11 and Figure 12 The diagram shows the internal flow after removing the solid structure. The turbine guide vane cooling structure also includes a rear chamber air inlet 32 ​​and a front chamber air inlet 31. The rear chamber air inlet 32 ​​is located at the upper end of the turbine guide vane and is connected to the rear chamber II. The front chamber air inlet 31 is located at the upper end of the turbine guide vane and is connected to the front chamber I.

[0072] The cooling process of the turbine guide vane is as follows.

[0073] The first stream of cold air L enters the impact bushing 2 from the front chamber inlet 31 at the lower end of the turbine guide vane 1. Part of the cold air impacts the inner side of the single-layer wall 12 through the bushing impact hole 21 on the impact bushing 2, and forms a transverse flow between the impact bushing 2 and the single-layer wall 12. It flows out of the guide vane through the air film hole on the surface of the single-layer wall 12, forming an air film protection on the outer side of the wall. The other part of the cold air enters the impact chamber III through the chamber impact hole 51 on the partition plate 5, impacts the inner wall at the back of the blade, and flows out of the guide vane through the impact chamber air film hole 52 on the surface.

[0074] The second stream of cold air G enters the rear cavity II from the rear cavity inlet 32 ​​at the upper end of the turbine guide vane 1. The cold air is subjected to impact cooling on the inner side of the outer wall 13 through the inner wall impact hole 17 on the inner wall 14, and forms a transverse flow in the double wall partition 16. At this time, the gap turbulence column 18 arranged between the double walls enhances the turbulence of the flow and strengthens the convective heat transfer effect of the cold air between the double walls. Through the double cooling structure on the suction side, the cold air flows to the blade trailing edge cavity IV. The trailing edge turbulence column 19 at the trailing edge cavity IV further enhances the convective heat transfer and finally flows out through the trailing edge slit 161. Another part of the cold air in the rear cavity II flows to the impact cavity III through the partition impact hole 22 on the partition plate 11, increasing the amount of cold air.

[0075] Therefore, the aforementioned turbine guide vane, by setting a hybrid structure with single-layer and double-layer walls, adopts a structure combining single-layer walls and irregularly shaped film cooling holes in the guide vane's front cavity, and further employs an impact chamber structure in the blade's back and axle area, with air supplied simultaneously from the front and rear cavities, thereby improving the cooling and heat transfer in the blade's back and axle area, and thus improving the film cooling coverage effect on the blade's back; the guide vane's rear cavity adopts a double-layer wall cooling design structure, which enhances the internal heat transfer characteristics of the wall surface through the double-layer wall structure, thereby improving the overall cooling effect of the blade.

[0076] Under certain cooling air volume constraints, based on the heat load distribution on the guide vane surface, the single-layer wall and irregular film cooling hole structure of the front cavity, the impact cavity structure of the blade back cheek area, and the double-layer wall cooling structure of the rear cavity are used to improve the film cooling protection of the local area on the blade back and enhance the wall heat transfer of the rear cavity, thereby improving the overall cooling effect of the guide vane.

[0077] Figure 13A ,13B Simulation diagrams of the film cooling effects near the throat of the turbine guide vane's suction surface in an aerodynamic environment are presented, showing the effects of a double-walled structure with circular film cooling holes and a single-walled structure with irregularly shaped film cooling holes. The horizontal and vertical axes represent the chordal and spanwise lengths of the blade. Simulation results show that, compared to... Figure 13A The double-walled and circular film pore structure shown is... Figure 13B The single-walled and irregularly shaped air film vents have a higher distribution range and value of air film cooling effect than the former, especially at the upper part of the air film vent array, where the cooling effect is particularly obvious.

[0078] The air film pores can be selected as follows: Figures 14A-14B The expansion-type air film orifice is shown. The orifice diameter c4 of the expansion-type air film orifice is including but not limited to 0.5mm to 0.7mm, the inclination angle θ4 is including but not limited to 30° to 40°, the expansion angle β4 is including but not limited to 10° to 15°, and the ratio of the length l4 of the straight section 142 to the length s4 of the expansion section 142 is including but not limited to (1.2 to 1.7):1.

[0079] Membrane pores can also be selected Figures 15A-15B The W-shaped air film orifice 151 shown has an orifice diameter c5 of, but is not limited to, 0.6 mm to 0.7 mm, an inclination angle θ5 of, but is not limited to, 30° to 40°, a back tilt angle ψ5 of, but is not limited to, 10° to 15°, an expansion angle β5 of, but is not limited to, 10° to 13°, a V-shaped included angle γ5 of, but is not limited to, 120° to 140°, and a ratio of the length l3 of the straight hole section 132 to the length s3 of the expansion section 133 of (1.2 to 1.7):1.

[0080] Figure 16 Another structure for the chambers in the turbine guide vane is shown. In this embodiment, the front chamber I is a single-layer wall and impact bushing structure, employing... Figure 17 and Figure 18 The diagram illustrates a combination of impact and film cooling; the suction surface 103 of the rear cavity II is configured as a combination of impact and turbulence column. Based on the design requirements of aero-engines, and taking into account the aerodynamic parameters of high-pressure turbines and the limitations of cooling gas usage, the geometric parameters such as the shape and quantity of the irregular film cooling holes in the guide vane, as well as the specific structural form of the double-walled cooling unit, can be adaptively adjusted, demonstrating good applicability.

[0081] In the description of this application, it should be understood that the orientation or positional relationship indicated by directional terms such as "front, back, up, down, left, right", "horizontal, vertical, horizontal" and "top, bottom" is usually based on the orientation or positional relationship shown in the accompanying drawings, and is only for the convenience of describing this application and simplifying the description. Unless otherwise stated, these directional terms do not indicate or imply that the device or element referred to must have a specific orientation or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation on the scope of protection of this application; the directional terms "inner" and "outer" refer to the inner and outer contours relative to the outline of each component itself.

[0082] Furthermore, this application uses specific terms to describe embodiments of the application. For example, "an embodiment," "one embodiment," and / or "some embodiments" refer to a particular feature, structure, or characteristic related to at least one embodiment of the application. Therefore, it should be emphasized and noted that "an embodiment," "one embodiment," or "an alternative embodiment" mentioned twice or more in different locations in this specification do not necessarily refer to the same embodiment. In addition, certain features, structures, or characteristics in one or more embodiments of the application can be appropriately combined.

[0083] While the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the invention. Any variations and modifications can be made by those skilled in the art without departing from the spirit and scope of the invention. Therefore, any modifications, equivalent changes, and alterations made to the above embodiments based on the technical essence of the present invention, without departing from the scope of the invention, fall within the protection scope defined by the claims of the present invention.

Claims

1. A turbine guide vane cooling structure, disposed on a turbine guide vane, the turbine guide vane comprising a suction-side blade wall and a pressure-side blade wall, and further comprising a front cavity and a rear cavity divided by a partition plate located inside the blade body, characterized in that, Also includes: The partition plate is connected to the cavity partition plate and the suction side blade wall on both sides, respectively, to divide the front cavity into an impact cavity. The impact cavity is located in the leaf back cheek area. The partition plate is provided with a partition plate impact hole to connect the impact cavity and the front cavity. The cavity partition plate is provided with a cavity impact hole to connect the impact cavity and the rear cavity. The front cavity is provided with a single-layer wall and an impact bushing, while the rear cavity is provided with a double-layer wall cooling structure.

2. The turbine guide vane cooling structure as described in claim 1, characterized in that, The impact bushing is provided with bushing impact holes, and the single-layer wall is provided with air film holes.

3. The turbine guide vane cooling structure as described in claim 2, characterized in that, At least some of the air film pores on the single-layer wall are irregularly shaped air film pores.

4. The turbine guide vane cooling structure as described in claim 1, characterized in that, The double-wall cooling structure includes a suction-side cooling structure and a pressure-side cooling structure.

5. The turbine guide vane cooling structure as described in claim 1, characterized in that, The pressure-side cooling structure includes an inner wall, an outer wall, and a gap turbulence column located between them. The inner wall and the outer wall are provided with air film pores, or include an inner wall, an outer wall, and air film pores located on the inner wall and the outer wall. The suction-side cooling structure includes an inner wall, an outer wall, and a gap turbulence column located between them. The inner wall and the outer wall are provided with air film pores, or include an inner wall, an outer wall, and air film pores located on the inner wall and the outer wall.

6. The turbine guide vane cooling structure as described in claim 5, characterized in that, The air film pores on the inner wall and the outer wall are one or more of the following: conical air film pores, scoop-shaped air film pores, expanded air film pores, and W-shaped air film pores.

7. The turbine guide vane cooling structure as described in claim 1, characterized in that, The turbine guide vane cooling structure also includes a tail section partition, on both sides of which are connected to the suction side blade wall and the pressure side blade wall, thereby dividing the rear cavity into a trailing edge cavity.

8. The turbine guide vane cooling structure as described in claim 7, characterized in that, The double-wall cooling structure includes an inner wall and an outer wall. The tail ends of the inner wall, located on the pressure side and the suction side, are connected to form the tail partition.

9. The turbine guide vane cooling structure as described in claim 7, characterized in that, The trailing edge cavity is provided with multiple rows of trailing edge baffles and trailing edge slits.

10. The turbine guide vane cooling structure as described in claim 9, characterized in that, The diameter of the trailing edge spoiler column ranges from 1 mm to 1.8 mm.

11. The turbine guide vane cooling structure as described in claim 1, characterized in that, The thickness of the single-layer wall ranges from 1.6 mm to 2.2 mm, and / or The wall thickness of the impact bushing ranges from 0.2 mm to 0.5 mm, and / or The impact distance between the impact bushing and the single-layer wall ranges from 0.4 mm to 1.0 mm.

12. The turbine guide vane cooling structure as described in claim 5, characterized in that, The wall thickness of the double-layer wall ranges from 2.1 mm to 2.7 mm, and / or The pore size of the air film pores on the inner wall ranges from 0.8 mm to 1.0 mm, and / or The diameter of the gap turbulence column ranges from 0.8 mm to 1.4 mm.

13. The turbine guide vane cooling structure as described in claim 1, characterized in that, The diameter of the cavity impact hole ranges from 1.2 mm to 1.4 mm.

14. The turbine guide vane cooling structure as described in claim 1, characterized in that, The suction side blade wall forming the impact cavity is provided with multiple exhaust film holes, and the outlet of the last exhaust film hole is located on the front side of the blade throat.

15. The turbine guide vane cooling structure as described in claim 1, characterized in that, The turbine guide vane cooling structure also includes a rear chamber air inlet and a front chamber air inlet. The rear chamber air inlet is located at the upper end of the turbine guide vane and communicates with the rear chamber. The front chamber air inlet is located at the upper end of the turbine guide vane and communicates with the front chamber.

16. A turbine guide vane, characterized in that, Includes the turbine guide vane cooling structure as described in any one of claims 1-15.

Images